Unitary flow path structure

ABSTRACT

Gas turbine engines, as well as outer walls and flow path assemblies of gas turbine engines, are provided. For example, an outer wall of a flow path comprises a combustor portion extending through a combustion section, and a turbine portion extending through at least a first turbine stage of a turbine section. The combustor and turbine portions are integrally formed as a single unitary structure that defines an outer boundary of the flow path. As another example, a flow path assembly comprises a combustor dome positioned a forward end of a combustor; an outer wall extending from the combustor dome through at least a first turbine stage; and an inner wall extending from the combustor dome through at least a combustion section. The combustor dome extends radially from the outer wall to the inner wall and is integrally formed with the outer and inner walls as a single unitary structure.

FIELD

The present subject matter relates generally to gas turbine engines.More particularly, the present subject matter relates to unitarystructures for defining a flow path within a gas turbine engine.

BACKGROUND

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine generally includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gasesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

More particularly, the combustion section includes a combustor having acombustion chamber defined by a combustor liner. Downstream of thecombustor, the turbine section includes one or more stages, for example,each stage may a plurality of stationary nozzle airfoils as well as aplurality of blade airfoils attached to a rotor that is driven by theflow of combustion gases against the blade airfoils. The turbine sectionmay have other configurations as well, e.g., the turbine may be acounter-rotating turbine without stationary nozzle airfoils. In anyevent, a flow path is defined by an inner boundary and an outerboundary, which both extend from the combustor through the stages of theturbine section.

Typically, the inner and outer boundaries defining the flow pathcomprise separate components. For example, an outer liner of thecombustor, a separate outer band of a nozzle portion of a turbine stage,and a separate shroud of a blade portion of the turbine stage usuallydefine at least a portion of the outer boundary of the flow path.Utilizing separate components to form each of the outer boundary and theinner boundary may require one or more seals at each interface betweenthe separate components to minimize leakage of fluid from the flow path.Thus, a great number of parts may be required to construct a flow pathassembly, which can increase the complexity and weight of the gasturbine engine. Further, although seals may be provided, utilizingseparate components in the flow path assembly provides several pointsfor leakage of the fluid from the flow path. Increased weight,complexity, and leakage can negatively impact engine performance, aswell as assembly of the engine during manufacturing.

Accordingly, improved flow path assemblies would be desirable. Forexample, a unitary outer boundary structure extending through thecombustion section and at least a first stage of the turbine sectionwould be beneficial. Further, a flow path assembly comprising a unitaryinner boundary structure and a unitary outer boundary structure would beuseful. Additionally, a flow path assembly comprising an integralcombustor dome, inner boundary structure, and outer boundary structurewould be helpful. Moreover, a gas turbine engine having a flow pathassembly with a unitary outer boundary structure would be advantageous.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, an outer wall ofa flow path of a gas turbine engine is provided. The outer wallcomprises a combustor portion extending through a combustion section ofthe gas turbine engine; and a turbine portion extending through at leasta first turbine stage of a turbine section of the gas turbine engine.The combustor portion and the turbine portion are integrally formed suchthat the combustor portion and the turbine portion are a single unitarystructure. The unitary structure defines an outer boundary of the flowpath.

In another exemplary embodiment of the present disclosure, a flow pathassembly for a gas turbine engine is provided. The flow path assemblycomprises a combustor dome positioned a forward end of a combustor of acombustion section of the gas turbine engine. The flow path assemblyfurther comprises an outer wall extending from the combustor domethrough at least a first turbine stage of a turbine section of the gasturbine engine. The flow path assembly also comprises an inner wallextending from the combustor dome through at least the combustionsection. The combustor dome extends radially from the outer wall to theinner wall. The combustor dome, outer wall, and inner wall areintegrally formed such that the combustor dome, outer wall, and innerwall are a single unitary structure.

In a further exemplary embodiment of the present disclosure, a gasturbine engine is provided. The gas turbine engine comprises acombustion section and a turbine section. The turbine section includes afirst turbine stage positioned immediately downstream of the combustionsection and a second turbine stage positioned immediately downstream ofthe first turbine stage. The combustion section and the turbine sectiondefine a flow path. The combustion section includes an outer linerdefining an outer boundary of the flow path through the combustionsection. Further, each of the first turbine stage and the second turbinestage includes a nozzle portion and a blade portion. Each nozzle portioncomprises an outer band defining an outer boundary of the flow paththrough the nozzle portion. Each blade portion comprises a shrouddefining an outer boundary of the flow path through the blade portion.The outer liner, the outer bands, and the shrouds are integrally formedsuch that the outer liner, the outer bands, and the shrouds are a singleunitary structure.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-section view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIG. 2 provides a schematic exploded cross-section view of a combustionsection and a high pressure turbine section of the gas turbine engine ofFIG. 1 according to an exemplary embodiment of the present subjectmatter.

FIG. 3A provides a schematic cross-section view of the combustionsection and high pressure turbine section of FIG. 2 according to anexemplary embodiment of the present subject matter.

FIGS. 3B, 3C, 3D, and 3E provide schematic cross-section views of thecombustion section and high pressure turbine section of FIG. 2 accordingto other exemplary embodiments of the present subject matter.

FIG. 3F provides a partial perspective view of a portion of an integralouter boundary structure and inner boundary structure of the combustionsection and high pressure turbine section of FIG. 2 according to anexemplary embodiment of the present subject matter.

FIGS. 4A, 4B, 4C, 5A, 5B, and 5C provide schematic cross-section viewsof the combustion section and high pressure turbine section of FIG. 2according to other exemplary embodiments of the present subject matter.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first,” “second,” and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. In other embodiments ofturbofan engine 10, additional spools may be provided such that engine10 may be described as a multi-spool engine.

For the depicted embodiment, fan section 14 includes a fan 38 having aplurality of fan blades 40 coupled to a disk 42 in a spaced apartmanner. As depicted, fan blades 40 extend outward from disk 42 generallyalong the radial direction R. The fan blades 40 and disk 42 are togetherrotatable about the longitudinal axis 12 by LP shaft 36. In someembodiments, a power gear box having a plurality of gears may beincluded for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

It will be appreciated that, although described with respect to turbofan10 having core turbine engine 16, the present subject matter may beapplicable to other types of turbomachinery. For example, the presentsubject matter may be suitable for use with or in turboprops,turboshafts, turbojets, industrial and marine gas turbine engines,and/or auxiliary power units.

In some embodiments, components of turbofan engine 10, particularlycomponents within hot gas path 78, such as components of combustionsection 26, HP turbine 28, and/or LP turbine 30, may comprise a ceramicmatrix composite (CMC) material, which is a non-metallic material havinghigh temperature capability. Of course, other components of turbofanengine 10, such as components of HP compressor 24, may comprise a CMCmaterial. Exemplary CMC materials utilized for such components mayinclude silicon carbide (SiC), silicon, silica, or alumina matrixmaterials and combinations thereof. Ceramic fibers may be embeddedwithin the matrix, such as oxidation stable reinforcing fibers includingmonofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6),as well as rovings and yarn including silicon carbide (e.g., NipponCarbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning'sSYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and choppedwhiskers and fibers (e.g., Nextel's 440 and SAFFIL®), and optionallyceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinationsthereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica,talc, kyanite, and montmorillonite). For example, in certainembodiments, bundles of the fibers, which may include a ceramicrefractory material coating, are formed as a reinforced tape, such as aunidirectional reinforced tape. A plurality of the tapes may be laid uptogether (e.g., as plies) to form a preform component. The bundles offibers may be impregnated with a slurry composition prior to forming thepreform or after formation of the preform. The preform may then undergothermal processing, such as a cure or burn-out to yield a high charresidue in the preform, and subsequent chemical processing, such asmelt-infiltration or chemical vapor infiltration with silicon, to arriveat a component formed of a CMC material having a desired chemicalcomposition. In other embodiments, the CMC material may be formed as,e.g., a carbon fiber cloth rather than as a tape.

As stated, components comprising a CMC material may be used within thehot gas path 78, such as within the combustion and/or turbine sectionsof engine 10. As an example, the combustion section 26 may include acombustor formed from a CMC material and/or one or more stages of one ormore stages of the HP turbine 28 may be formed from a CMC material.However, CMC components may be used in other sections as well, such asthe compressor and/or fan sections. Of course, in some embodiments,other high temperature materials and/or other composite materials may beused to form one or more components of engine 10.

FIG. 2 provides an exploded view of a schematic cross-section of thecombustion section 26 and the HP turbine 28 of the turbine section ofthe turbofan engine 10 according to an exemplary embodiment of thepresent subject matter. FIG. 3A provides an unexploded schematiccross-sectional view of the combustion section 26 and the HP turbine 28of FIG. 2 that focuses on an outer boundary of a flow path through thecombustion section 26 and HP turbine 28. The depicted combustion section26 includes a generally annular combustor 80, and downstream of thecombustion section 26, the HP turbine 28 includes a plurality of turbinestages. More particularly, for the depicted embodiment, HP turbine 28includes a first turbine stage 82 and a second turbine stage 84. Inother embodiments, the HP turbine 28 may comprise a different number ofturbine stages; for example, the HP turbine 28 may include one turbinestage or more than two turbine stages. The first turbine stage 82 ispositioned immediately downstream of the combustion section 26, and thesecond turbine stage 84 is positioned immediately downstream of thefirst turbine stage 82. Further, each turbine stage 82, 84 comprises anozzle portion and a blade portion; the first turbine stage 82 includesnozzle portion 82N and blade portion 82B, and the second turbine stage84 includes nozzle portion 84N and blade portion 84B. The nozzle portion82N of the first turbine stage 82 is located immediately downstream ofthe combustion section 26, such that the nozzle portion 82N of the firstturbine stage 82 also may be referred to as a combustor dischargenozzle. Moreover, combustor 80 defines a generally annular combustionchamber 86 such that the combustor 80 may be described as a generallyannular combustor.

Additionally, as described in greater detail below, a flow path 100through the combustion section 26 and the HP turbine 28 is defined by anouter boundary and an inner boundary of a flow path assembly 101. Theouter and inner boundaries form a flow path for the combustion gases 66through the combustion section 26 and HP turbine 28; thus, the flow path100 may comprise at least a portion of the hot gas path 78 describedabove. Further, in other embodiments, the flow path 100 also may extendthrough LP turbine 30 and jet exhaust 32; in still other embodiments,the flow path 100 also may extend forward upstream of the combustionsection 26, e.g., into HP compressor 24. As such, it will be appreciatedthat the discussion herein of the present subject matter with respect tocombustion section 26 and HP turbine 28 is by way of example only andalso may apply to different configurations of gas turbine engines andflow paths 100.

As shown in the exploded view of FIG. 2, the outer and inner boundariesmay be defined by an outer wall 102 and an inner wall 120, respectively,which may include several portions of the combustion section 26 and HPturbine 28. For instance, the combustor 80 includes an outer liner 108defining an outer boundary of the flow path through the combustor 80.Each nozzle portion 82N, 84N comprises an outer band defining an outerboundary of a flow path through the nozzle portion of each turbinestage, and each blade portion 82B, 84B comprises a shroud defining anouter boundary of a flow path through the blade portion of each turbinestage. More particularly, as shown in FIG. 2, the first turbine stagenozzle portion 82N comprises outer band 110, first turbine stage bladeportion 82B comprises shroud 112, second turbine stage nozzle portion84N comprises outer band 114, and second turbine stage blade portion 84Bcomprises shroud 116. These portions of the combustion section 26 and HPturbine 28 may comprise at least a portion of the outer wall 102, asdescribed in greater detail below.

Further, as illustrated in FIG. 2, the combustor 80 includes an innerliner 122 defining an inner boundary of the flow path through thecombustor 80. Each nozzle portion 82N, 84N comprises an inner banddefining an inner boundary of the flow path through the nozzle portionof each turbine stage, and each blade portion 82B, 84B comprises one ormore blade platforms that define an inner boundary of the flow paththrough the blade portion of each turbine stage. More particularly, asshown in FIG. 2, the first turbine stage nozzle portion 82N comprisesinner band 124, first turbine stage blade portion 82B comprises bladeplatforms 132, second turbine stage nozzle portion 84N comprises innerband 136, and second turbine stage blade portion 84B comprises bladeplatforms 132. These portions of the combustion section 26 and HPturbine 28 may comprise at least a portion of the inner wall 122, asdescribed in greater detail below.

Moreover, in the depicted embodiment, a combustor dome 118 extendsradially across a forward end 88 of the combustor 80. The combustor dome118 may be a part of outer wall 102, may be a part of inner wall 120,may be a part of both outer wall 102 and inner wall 120 (e.g., a portionof the combustor dome 118 may be defined by the outer wall 102 and theremainder may be defined by the inner wall 120), or may be a separatecomponent from outer wall 102 and inner wall 120. Additionally, aplurality of nozzle airfoils is positioned in each of the nozzleportions 82N, 84N. Each nozzle airfoil 126 within the first turbinestage nozzle portion 82N extends radially from the outer band 110 to theinner band 124, and the nozzle airfoils 126 are spaced circumferentiallyabout the longitudinal centerline 12. Each nozzle airfoil 128 within thesecond turbine stage nozzle portion 84N extends radially from the outerband 114 to the inner band 136, and the nozzle airfoils 128 are spacedcircumferentially about the longitudinal centerline 12. Further, aplurality of blade airfoils 130 are positioned in each of the bladeportions 82B, 84B. Each blade airfoil 130 within the first turbine stageblade portion 82B is attached to blade platform 132, which in turn isattached to a first stage rotor 134. The blade airfoils 130 attached tothe first stage rotor 134 are spaced circumferentially about thelongitudinal centerline 12. Similarly, each blade airfoil 130 within thesecond turbine stage blade portion 84B is attached to a blade platform132, which in turn is attached to a second stage rotor 138. The bladeairfoils 130 attached to the second stage rotor 138 are spacedcircumferentially about the longitudinal centerline 12. Each bladeairfoils 130 extends radially outward toward the outer wall 102, i.e.,the outer boundary of the flow path 100, and a clearance gap is definedbetween a tip 140 of each blade airfoil 130 and the outer wall 102 suchthat each turbine rotor 134, 138 is free to rotate within its respectiveturbine stage. Although not depicted, each turbine rotor 134, 138 of theHP turbine 28 is connected to the HP shaft 34 (FIG. 1). In such manner,rotor blade airfoils 130 may extract kinetic energy from the flow ofcombustion gases through the flow path 100 defined by the HP turbine 28as rotational energy applied to the HP shaft 34.

Accordingly, flow path 100 through the combustion section 26 and the HPturbine 28 is defined by a flow path assembly 101 having an innerboundary and an outer boundary, and the inner and outer boundariesdefine the flow path for the combustion gases 66 through the combustionsection 26 and HP turbine 28. Portions of the outer boundary of the flowpath assembly 101 may be integrated or unified into a single piece outerwall 102 that defines the radially outer boundary of the gas flow path100. For instance, the outer wall 102 may include a combustor portion104 extending through a combustion section, such as combustion section26, and a turbine portion 106 extending through at least a first turbinestage of a turbine section, such as first turbine stage 82 of HP turbine28. The combustor portion 104 and turbine portion 106 are integrallyformed such that the combustor portion and the turbine portion are asingle unitary structure, i.e., a unitary outer wall 102.

In the exemplary embodiment depicted in FIG. 3A, the outer wall 102includes a combustor portion 104 extending through the combustionsection 26 and a turbine portion 106 extending through at least thefirst turbine stage 82 and the second turbine stage 84 of the turbinesection. In other embodiments, the turbine portion 106 may extendthrough fewer stages (e.g., through one turbine stage as just described)or through more stages (e.g., through one or more stages of the LPturbine 30 positioned downstream of HP turbine 28). The combustorportion 104 and the turbine portion 106 are integrally formed such thatthe combustor portion 104 and the turbine portion 106 are a singleunitary structure, which is referred to herein as unitary outer wall102.

The term “unitary” as used herein denotes that the associated component,such as the outer wall 102, is made as a single piece duringmanufacturing, i.e., the final unitary component is a single piece.Thus, a unitary component has a construction in which the integratedportions are inseparable and is different from a component comprising aplurality of separate component pieces that have been joined togetherand, once joined, are referred to as a single component even though thecomponent pieces remain distinct and the single component is notinseparable (i.e., the pieces may be re-separated). The final unitarycomponent may comprise a substantially continuous piece of material, orin other embodiments, may comprise a plurality of portions that arepermanently bonded to one another. In any event, the various portionsforming a unitary component are integrated with one another such thatthe unitary component is a single piece with inseparable portions.

As shown in FIG. 3A, the combustor portion 104 of the unitary structureforming outer wall 102 includes the outer liner 108 of the combustor 80.The turbine portion 106 includes the outer band 110 of the first turbinestage nozzle portion 82N, the shroud 112 of the first turbine stageblade portion 82B, the outer band 114 of the second turbine stage nozzleportion 84N, and the shroud 116 of the second turbine stage bladeportion 84B. As stated, these outer boundary components are integratedinto a single piece to form the unitary structure that is outer wall102. Thus, in the exemplary embodiment of FIG. 2, outer liner 108, outerband 110, shroud 112, outer band 114, and shroud 116 are integrallyformed, i.e., constructed as a single unit or piece to form theintegrated or unitary outer wall 102.

In some embodiments, other portions of the flow path assembly 101 may beintegrated into the unitary structure of outer wall 102, and in stillother embodiments, at least a portion of the outer boundary and theinner boundary are made as a single, unitary component such that theflow path assembly 101 may be referred to as an integrated flow pathassembly. For example, referring to FIG. 3B, the combustor portion 104of unitary outer wall 102 also may include the combustor dome 118 thatextends across the forward end 88 of combustor 80. As such, in theexemplary embodiment of FIG. 3B, the outer liner 108, outer band 110,shroud 112, outer band 114, shroud 116, and combustor dome 118 areconstructed as a single unit or piece to form the integrated or unitaryouter wall 102. That is, the outer liner 108, outer bands 110, 114,shrouds 112, 116, and combustor dome 118 are integrally formed such thatthe outer liner 108, outer bands 110, 114, shrouds 112, 116, andcombustor dome 118 are a single unitary structure.

As another example, referring to FIG. 3C, at least a portion of theinner wall 120 defining the inner boundary of the flow path 100 may beintegrated with the outer wall 102 to form an integrated flow pathassembly 101. In the exemplary embodiment of FIG. 3C, the combustorportion 104 further comprises the inner liner 122, such that the innerliner 122 is integrated with the unitary structure of the outer wall 102shown in FIG. 3B. Thus, the outer liner 108, outer band 110, shroud 112,outer band 114, shroud 116, combustor dome 118, and inner liner 122 areintegrally formed such that the outer liner 108, outer bands 110, 114,shrouds 112, 116, combustor dome 118, and inner liner 122 are a singleunitary structure. In the exemplary embodiment of FIG. 3D, the turbineportion 106 further includes the inner band 124 of the first turbinestage nozzle portion 82N, such that the inner band 124 is integratedwith the unitary structure of the flow path assembly 101 shown in FIG.3C. Accordingly, the outer liner 108, outer band 110, shroud 112, outerband 114, shroud 116, combustor dome 118, inner liner 122, and innerband 124 are integrally formed such that the outer liner 108, outerbands 110, 114, shrouds 112, 116, combustor dome 118, inner liner 122,and inner band 124 are a single unitary structure. In the exemplaryembodiment of FIG. 3E, the turbine portion 106 further includes theplurality of nozzle airfoils 126, such that each nozzle airfoil 126 ofthe plurality of nozzle airfoils 126 of the first turbine stage nozzleportion 82N is integrated with the unitary structure of the flow pathassembly 101 shown in FIG. 3D. Therefore, the outer liner 108, outerband 110, shroud 112, outer band 114, shroud 116, combustor dome 118,inner liner 122, inner band 124, and nozzle airfoils 126 are integrallyformed such that the outer liner 108, outer bands 110, 114, shrouds 112,116, combustor dome 118, inner liner 122, inner band 124, and nozzleairfoils 126 are a single unitary structure.

Of course, the nozzle airfoils 126 of the first turbine stage nozzleportion 82N may be integrated with the outer wall 102 without beingintegrated with the inner wall 120. For example, the plurality of nozzleairfoils 126 may be formed as a single unit or piece with the outerliner 108, outer band 110, shroud 112, outer band 114, shroud 116 suchthat the outer liner 108, outer bands 110, 114, shrouds 112, 116, andnozzle airfoils 126 are a single unitary structure, i.e., a unitaryouter wall 102. In other embodiments, the unitary outer wall 102 alsomay include the combustor dome 118, such that the outer liner 108, outerband 110, shroud 112, outer band 114, shroud 116, combustor dome 118,and nozzle airfoils 126 are integrally formed or constructed as a singleunit or piece. In still other embodiments, the inner liner 122 also maybe included, such that the outer liner 108, outer band 110, shroud 112,outer band 114, shroud 116, combustor dome 118, inner liner 122, andnozzle airfoils 126 are integrally formed as a single unitary structure,i.e., a unitary outer wall 102.

FIG. 3F provides a partial perspective view of a portion of an integralflow path assembly 101, having an outer wall 102 and inner wall 120formed as a single piece component. As described with respect to FIG. 3Dand shown in FIG. 3F, in some embodiments of the combustion gas flowpath assembly 101, the outer liner 108, outer band 110, shroud 112,outer band 114, shroud 116, combustor dome 118, inner liner 122, andinner band 124 are integrally formed such that the outer liner 108,outer bands 110, 114, shrouds 112, 116, combustor dome 118, inner liner122, and inner band 124 are a single unitary structure. FIG. 3F furtherillustrates that a plurality of openings 142 for receipt of fuel nozzleassemblies 90 and/or swirlers 92 may be defined in the forward end 88 ofcombustor 80 of the unitary flow path assembly 101. Further, it will beappreciated that FIG. 3F illustrates only a portion of the integral flowpath assembly 101 and that, although its entire circumference is notillustrated in FIG. 3F, the flow path assembly 101 is a single, unitarypiece circumferentially as well as axially. As such, the integral flowpath assembly 101 defines a generally annular, i.e., generallyring-shaped, flow path between the outer wall 102 and inner wall 120.

Integrating various components of the outer and inner boundaries of theflow path assembly 101 as described above can reduce the number ofseparate pieces or components within engine 10, as well as reduce theweight, leakage, and complexity of the engine 10, compared to known gasturbine engines. For instance, known gas turbine engines employ seals orsealing mechanisms at the interfaces between separate pieces of the flowpath assembly to attempt to minimize leakage of combustion gases fromthe flow path. By integrating the outer boundary, for example, asdescribed with respect to unitary outer wall 102, split points orinterfaces between the outer combustor liner and first turbine stageouter band, the first turbine stage outer band and the first turbinestage shroud, etc. can be eliminated, thereby eliminating leakage pointsas well as seals or sealing mechanisms required to prevent leakage.Similarly, by integrating components of the inner boundary, split pointsor interfaces between the integrated inner boundary components areeliminated, thereby eliminating leakage points and seals or sealingmechanisms required at the inner boundary. Accordingly, undesiredleakage, as well as unnecessary weight and complexity, can be avoided byutilizing unitary components in the flow path assembly. Other advantagesof unitary outer wall 102, unitary inner wall 120, and/or a unitary flowpath assembly 101 will be appreciated by those of ordinary skill in theart.

As illustrated in FIGS. 3A through 3F, the outer wall 102 and the innerwall 120 define a generally annular flow path therebetween. That is, theunitary outer wall 102 circumferentially surrounds the inner wall 120;stated differently, the unitary outer wall 102 is a single pieceextending 360° degrees about the inner wall 120, thereby defining agenerally annular or ring-shaped flow path therebetween. As such, thecombustor dome 118, which extends across the forward end 88 of thecombustor 80, is a generally annular combustor dome 118. Further, thecombustor dome 118 defines an opening 142 for receipt of a fuel nozzleassembly 90 positioned at forward end 88. The fuel nozzle assembly 90,e.g., provides combustion chamber 86 with a mixture of fuel andcompressed air from the compressor section, which is combusted withinthe combustion chamber 86 to generate a flow of combustion gases throughthe flow path 100. The fuel nozzle assembly 90 may attach to thecombustor dome 118 or may “float” relative to the combustor dome 118 andthe flow path 100, i.e., the fuel nozzle assembly 90 may not be attachedto the combustor dome 118. In the illustrated embodiments, the fuelnozzle assembly 90 includes a swirler 92, and in some embodiments, theswirler 92 may attach to the combustor dome 118, but alternatively, theswirler 92 may float relative to the combustor dome 118 and flow path100. It will be appreciated that the fuel nozzle assembly 90 or swirler92 may float relative to the combustor dome 118 and flow path 100 alongboth a radial direction R and an axial direction A or only along one orthe other of the radial and axial directions R, A. Further, it will beunderstood that the combustor dome 118 may define a plurality ofopenings 142, each opening receiving a swirler 92 or other portion offuel nozzle assembly 90.

As further illustrated in FIGS. 3A through 3F, as well as FIGS. 4Athrough 4C and FIGS. 5A and 5B discussed in greater detail below, theflow path assembly 101 generally defines a converging-diverging flowpath 100. More particularly, the outer wall 102 and the inner wall 120define a generally annular combustion chamber 86, which forms a forwardportion of the flow path 100. Moving aft or downstream of combustionchamber 86, the outer wall 102 and inner wall 120 converge toward oneanother, generally in the region of first turbine stage 82. Continuingdownstream of the first turbine stage 82, the outer wall 102 and innerwall 120 then diverge, generally in the region of second turbine stage84. The outer wall 102 and inner wall 120 may continue to divergedownstream of the second turbine stage 84. In exemplary embodiments,e.g., as shown in FIG. 3A and referring only to the unitary outer wall102, the first turbine stage nozzle outer band portion 110 and bladeshroud portion 112 of the outer wall 102 converge toward the axialcenterline 12. The second turbine stage nozzle outer band portion 114and blade shroud portion 116 of the outer wall 102 diverge away from theaxial centerline 12. As such, the outer boundary of flow path 100 formedby the unitary outer wall 102 defines a converging-diverging flow path100.

Turning to FIGS. 4A and 4B, other exemplary embodiments of the presentsubject matter are illustrated. FIG. 4A provides a schematiccross-sectional view of the combustion section 26 and the HP turbine 28of the turbine section according to one exemplary embodiment. FIG. 4Bprovides a schematic cross-sectional view of the combustion section 26and the HP turbine 28 of the turbine section according to anotherexemplary embodiment. FIG. 4C provides a schematic cross-sectional viewof the combustion section 26 and the HP turbine 28 of the turbinesection according to yet another exemplary embodiment.

In the embodiments shown in FIGS. 4A, 4B, and 4C, the outer wall 102 isformed as a single unitary structure and the inner wall 120 is formed asanother single unitary structure, and together, the unitary outer wall102 and the unitary inner wall 120 define the flow path 100. However, itshould be appreciated that the inner wall 120 need not be a singleunitary structure. For example, in the embodiments shown in FIGS. 4A,4B, and 4C, the inner wall 120 could comprise an inner liner 122 formedseparately from inner band 124.

As described with respect to FIGS. 3A through 3F, the unitary outer wall102 of FIGS. 4A, 4B, and 4C defines an outer boundary and the inner wall120 defines an inner boundary of the flow path 100. Together, theunitary outer wall 102 and the inner wall 120 form a flow path assembly101. The unitary outer wall 102 extends from the forward end 88 ofcombustor 80 of the combustion section 26 through at least the firstturbine stage 82 of the HP turbine 28, and in the depicted embodiments,the unitary outer wall 102 extends from forward end 88 to an aft end ofthe second turbine stage 84 of HP turbine 28. The inner wall 120includes at least the inner liner 122, and in embodiments in which theinner wall 120 is a unitary inner wall, the unitary inner wall 120extends from the forward end 88 of the combustor 80 through the firstturbine stage nozzle portion 82N. Accordingly, as shown in FIGS. 4A, 4B,and 4C, the outer wall 102 and inner wall 120 define the combustionchamber 86 of the combustor 80.

Like the embodiments described with respect to FIGS. 3A through 3F, theunitary outer wall 102 of the embodiments shown in FIGS. 4A, 4B, and 4Cincludes the outer liner 108, outer band 110, shroud 112, outer band114, and shroud 116. Further, in the exemplary embodiment of FIG. 4A,the unitary outer wall 102 includes the combustor dome 118 defined atthe forward end 88 of the combustor 80. Thus, the outer liner 108, outerbands 110, 114, shrouds 112, 116, and combustor dome 118 are integrallyformed or constructed as a single unitary structure, i.e., outer wall102 is a single unit or piece that includes combustor dome 118.Alternatively, as shown in the exemplary embodiment of FIG. 4B, theunitary outer wall 102 includes a radially outer portion of thecombustor dome 118, such that the outer liner 108, outer band 110,shroud 112, outer band 114, shroud 116, and a portion of the combustordome 118 are integrally formed or constructed as a single unitarystructure, i.e., outer wall 102 is a single unit or piece that includesa portion combustor dome 118.

Moreover, like the embodiments described with respect to FIGS. 3Athrough 3F, the inner wall 120 of the embodiments shown in FIGS. 4A, 4B,and 4C at least includes the inner liner 122 of the combustor 80. Insome embodiments, such as illustrated in FIGS. 4A and 4B, the inner wall120 also includes the inner band 124 of the first turbine stage nozzleportion 82N. In such embodiments, the inner liner 122 and inner band 124are integrally formed as a single unitary structure, i.e., as a singleunit or piece that may be referred to as unitary inner wall 120. Inother embodiments, as illustrated in FIG. 4B, the unitary inner wall 120may include a radially inner portion of the combustor dome 118 such thatthe inner liner 122 and the portion of the combustor dome 118 areintegrally formed or constructed as a single unitary structure or suchthat the inner liner 122, inner band 124, and the portion of thecombustor dome 118 are integrally formed or constructed as a singleunitary structure. That is, in some embodiments, the unitary inner wall120 is a single unit or piece that includes a portion of the combustordome 118 (and may or may not include the inner band 124). In still otherembodiments, as shown in FIG. 4C, the unitary inner wall 120 includesthe combustor dome 118 defined at the forward end 88 of the combustor80. Thus, the combustor dome 118 and inner liner 122 (as well as innerband 124 in some embodiments) are integrally formed or constructed as asingle unitary structure, i.e., inner wall 102 is a single unit or piecethat includes combustor dome 118.

Further, the first turbine stage nozzle airfoils 126 may be integratedwith the outer wall 102 and/or with the inner wall 120. As previouslydescribed, the first turbine stage nozzle airfoils 126 may be integratedwith the outer wall 102, but in other embodiments, the first turbinestage nozzle airfoils 126 may be integrated with the inner wall 120 andnot the outer wall 102 or may be integrated with both the outer andinner walls 102, 120. Whether formed separately from the walls 102, 120,integrated with the inner wall 120 to form a single unitary structurewith the inner wall 120, integrated with the outer wall 102 to form asingle unitary structure with the outer wall 102, or integrated withboth the outer and inner walls 102, 120 to form a single unitarystructure with the outer and inner walls 102, 120, a plurality of nozzleairfoils 126 extend from the inner wall 120 to the outer wall 102 withinthe first turbine stage nozzle portion 82N. Additionally, as describedabove, the first turbine stage 82 includes a first stage rotor 134having a plurality of rotor blade airfoils 130 attached thereto.Downstream of the first turbine stage 82, a plurality of nozzle airfoils128 extend from the inner band 136 to the outer wall 102 within thesecond turbine stage nozzle portion 84N, and the second turbine stageblade portion 84B includes a second stage rotor 138 having a pluralityof rotor blade airfoils 130 attached thereto.

In the embodiments of FIGS. 4A, 4B and 4C, the integrated or unitaryouter wall 102 extends circumferentially about the integrated or unitaryinner wall 120. That is, the outer wall 102 circumferentially surroundsthe inner wall 120 or the unitary outer wall 102 is a single pieceextending 360° degrees about the inner wall 120. As such, the outer wall102 and the inner wall 120 define a generally annular flow paththerebetween. Further, the combustor dome 118 extends across the forwardend 88 of the combustor 80, and whether integrated into the unitaryouter wall 102 in whole or in part or integrated into the unitary innerwall 120 in whole or in part, the combustor dome 118 is a generallyannular combustor dome 118.

In addition, the flow path assembly 101 illustrated in the embodimentsof FIGS. 4A, 4B and 4C includes at least one opening 142 for receipt ofa fuel nozzle assembly 90. As described with respect to FIGS. 3A through3F, in some embodiments, the fuel nozzle assembly 90 may attach to thecombustor dome 118, which may be integrated with the outer wall 102 inwhole as in the embodiment of FIG. 4A or in part as shown in FIG. 4B,where the remainder is integrated with the inner wall 120. As alsodescribed, the combustor dome 118 may be integrated with the inner wall120 in whole as illustrated in FIG. 4C, such that the fuel nozzleassembly 90 may attach to the combustor dome portion of unitary innerwall 120. In other embodiments, the fuel nozzle assembly 90 does notattach to the combustor dome 118 but floats relative to the combustordome 118 and the flow path 100. As depicted, the fuel nozzle assembly 90includes swirler 92, which may be the portion of fuel nozzle assembly 90that attaches to the combustor dome 118 or the portion that floatsrelative to the combustor dome 118 and flow path 100. As previouslydescribed, the fuel nozzle assembly 90 or swirler 92 may float relativeto the combustor dome 118 and flow path 100 along both the radialdirection R and the axial direction A or only along one or the other ofthe radial and axial directions R, A. Moreover, as shown in FIG. 3F, thecombustor dome 118 may define a plurality of openings 142, and eachopening may receive a swirler 92 or other portion of fuel nozzleassembly 90.

Referring still to FIGS. 4A, 4B, and 4C, the unitary outer wall 102 andthe inner wall 120 may define one or more features where the walls 102,120 meet up with one another and, in some embodiments, may be attachedto one another. For instance, in the embodiment of FIG. 4A, the outerwall 102 defines a flange 144 along a radially inner edge of the outerwall 102 at the forward end 88 of the combustor 80, and the inner wall120 defines a flange 146 along a forward edge at the combustor forwardend 88. In the embodiment of FIG. 4B, the outer wall flange 144 isdefined along an edge of the combustor dome portion of the unitary outerwall 102, and similarly, the inner wall flange 146 is defined along anedge of the combustor dome portion of the unitary inner wall 120. Asshown in FIG. 4C, the outer wall 102 may define the outer wall flange144 along a forward edge of the outer wall 102, and the inner wall 120,which includes combustor dome 118 in the illustrated embodiment, maydefine the inner wall flange 146 along a radially outer edge of theinner wall 120. FIGS. 4A, 4B, and 4C illustrate that the flow path 100may be discontinuous between the inner wall 120 and the outer wall 102,i.e., formed from a separate inner and outer boundaries rather thanintegral inner and outer boundaries as shown in FIGS. 3C through 3F.More particularly, the flow path 100 may be discontinuous where theouter wall flange 144 and the inner wall flange 146 are defined.

Thus, in the embodiment of FIG. 4A, the outer wall 102 may be secured tothe inner wall 120 at flanges 144, 146 near a radially inner, forwardportion of the combustor 80. Alternatively, the flanges 144, 146 asshown in FIG. 4A may define an area where the walls 102, 120 align ormeet up with one another, e.g., flanges 144, 146 may define a slip jointbetween walls 102, 120. In the embodiment of FIG. 4B, the outer wall 102may be secured to the inner wall 120 at flanges 144, 146 near a radialcenterline of the combustor dome 118. In other embodiments, the flanges144, 146 as illustrated in FIG. 4B may define an area where the walls102, 120 align or meet up with one another, e.g., flanges 144, 146 maydefine a slip joint between walls 102, 120. In alternative embodiments,such the embodiment of FIG. 4C, the outer wall 120 may be secured to theinner wall 120 at flanges 144, 146 near a radially outer, forwardportion of the combustor 80, or the flanges 144, 146 as shown in FIG. 4Cmay define an area where the walls 102, 120 align or meet up with oneanother, e.g., flanges 144, 146 may define a slip joint between walls102, 120 at a radially outer, forward portion of combustor 80. In stillother embodiments, the flanges 144, 146 may be defined in otherlocations such that the outer wall 102 and inner wall 120 are securedto, align, or meet up with one another at a location different fromthose depicted in FIGS. 4A, 4B, and 4C.

Any suitable fastener or other attachment means may be used to securethe outer and inner walls 102, 120 at the flanges 144, 146. For example,a plurality of apertures may be defined in each flange 144, 146, andeach aperture of the outer wall flange 144 may align with an aperture ofthe inner wall flange 146 for receipt of a fastener in each pair ofaligned apertures. It will be appreciated that the outer wall 102 andthe inner wall 120 may be attached to one another in other ways as well.Of course, in other embodiments as described above, the outer wall 102and inner wall 120 may not be secured to one another but may moveradially and/or axially with respect to one another.

Turning now to FIGS. 5A, 5B, and 5C, schematic cross-sectional views areprovided of the combustion section 26 and the HP turbine 28 of theturbine section of turbofan engine 10 according to other exemplaryembodiments of the present subject matter. Unlike the embodiments ofFIGS. 3B through 3F and FIGS. 4A through 4C, the combustor dome 118 ofthe embodiments shown in FIGS. 5A, 5B, and 5C is not integrated witheither the outer wall 102 or the inner wall 120 in whole or in part.That is, the combustor dome 118 is a separate component from both theouter wall 102 and the inner wall 120.

Accordingly, as shown in FIGS. 5A, 5B, and 5C, the outer wall 102 is aunitary outer wall including a combustor portion 104, which extendsthrough the combustion section 26 of engine 10, and a turbine portion106, which extends through at least a first turbine stage of the turbinesection of engine 10. In the embodiments shown in FIGS. 5A through 5C,the unitary outer wall 102 extends through the combustion section 26 toan aft end of HP turbine 28, which includes two turbine stages 82, 84.The combustor portion 104 and turbine portion 106 are integrally formedas a single unitary structure, i.e., unitary outer wall 102. Forexample, as shown and described with respect to FIG. 3A, the combustorportion 104 of the unitary outer wall 102 comprises the outer liner 108of combustor 80. The turbine portion 106 of unitary outer wall 102comprises outer band 110 of first turbine stage nozzle portion 82N, theshroud 112 of the first turbine stage blade portion 82B, the outer band114 of the second turbine stage nozzle portion 84N, and the shroud 116of the second turbine stage blade portion 84B. The turbine portion 106of unitary outer wall 102 also may include a plurality of nozzleairfoils 126, which are integrally formed or constructed with the outerliner 108, outer bands 110, 114, and shrouds 112, 116 to form a singleunitary structure, i.e., as a single unit or piece.

Further, as depicted in FIGS. 5A, 5B, and 5C, the inner wall 120 extendsfrom the forward end 88 of the combustor 80 through at least thecombustion section 26. For instance, the inner wall 120 may compriseseparate components defining the inner boundary of the flow path 100. Inother embodiments, the inner wall 120 may be a unitary inner wall 120including an inner liner 122 and inner band 124 integrally formed as asingle unitary structure, i.e., as a single unit or piece. As anotherexample, the inner wall 120 may be a unitary inner wall 120 includinginner liner 122, inner band 124, and first turbine stage nozzle airfoils126 integrally formed as a single unitary structure, i.e., as a singleunit or piece. Further, in the depicted embodiments of FIGS. 5A, 5B, and5C, the flow path 100 may be discontinuous between the inner wall 120and the outer wall 102, i.e., formed from a separate inner and outerboundaries rather than integral inner and outer boundaries as shown inFIGS. 3C through 3F. More particularly, the flow path 100 may bediscontinuous between the combustor dome 118 and outer wall 102, as wellas between combustor dome 118 and inner wall 120.

Referring particularly to FIG. 5A, the combustor dome 118 is positionedat forward end 88 of combustor 80 of combustion section 26 and extendsradially from the outer wall 102 to the inner wall 120. The combustordome 118 is configured to move axially with respect to the inner wall120 and the outer wall 102 but may be attached to, and accordinglysupported by, one or more fuel nozzle assemblies 90. More particularly,an axial slip joint 150 is formed between the combustor dome 118 andeach of the outer wall 102 and the inner wall 120 such that thecombustor dome 118 may move or float axially with respect to the innerwall 120 and outer wall 102. Allowing the combustor dome 118 to floatrelative to the outer wall 102 and inner wall 120 can help control theposition of the fuel nozzle assembly 90 with respect to the combustordome 118 and combustor 80. For example, the combustor dome 118, outerwall 102, and inner wall 120 may be made of a different material ormaterials than the fuel nozzle assembly 90. As described in greaterdetail below, in an exemplary embodiment, the combustor dome 118, outerwall 102, and inner wall 120 are made from a ceramic matrix composite(CMC) material, and the fuel nozzle assembly 90 may be made from ametallic material, e.g., a metal alloy or the like. In such embodiment,the CMC material thermally grows or expands at a different rate than themetallic material. Thus, allowing the combustor dome 118 to move axiallywith respect to outer and inner walls 102, 120 may allow for tightercontrol of the immersion of swirler 92 of fuel nozzle assembly 90 withincombustor dome 118, as well as combustor 80, than if the combustor dome118 was attached to the outer and inner walls 102, 120. Tighter controlof the position of fuel nozzle assembly 90 and its components withrespect to combustor 80 can reduce variation in operability andperformance of engine 10.

Further, the outer wall 102 and inner wall 120 also may move axially andradially with respect to the combustor dome 118. By decoupling thecombustor dome 118 from the walls 102, 120 and allowing relativemovement between the walls 102, 120 and the combustor dome 118, stresscoupling may be alleviated between the outer and inner walls 102, 120and the combustor dome 118. Moreover, any leakage between the uncoupledcombustor dome 118 and outer and inner walls 102, 120 may be utilized aspurge and/or film starter flow.

As illustrated in FIG. 5A, the combustor dome 118 includes an outer wing152 and an inner wing 154. The outer wing 152 extends aft along theouter wall 102, and the inner wing 154 extends aft along the inner wall120. The wings 152, 154 may help guide the combustor dome 118 as itmoves with respect to the outer wall 102 and inner wall 120, and thewings 152, 154 also may help maintain the radial position or alignmentof the combustor dome 118 as it moves axially. The wings may provide aconsistent gap between the dome 118 and walls 102, 120 for purge and/orfilm starter flow as previously described.

Turning to FIG. 5B, in other embodiments, each wing 152, 154 may extendforward from the combustor dome body 156, rather than aft as shown inFIG. 5A. The forward-extending wings 152, 154 may be used to mount thecombustor dome 118 to a component other than the fuel nozzle assembly90/swirler 92, e.g., to a metal dome supporting fuel nozzle assembly 90and/or to either or both of the outer wall 102 and inner wall 120 at theforward end 88 of combustor 80. In some embodiments, theforward-extending wings 152, 154 of combustor dome 118 may be pinned orotherwise attached to the outer wall 102 and the inner wall 120 as shownin FIG. 5B. In still other embodiments, one of the wings 152, 154 mayextend forward and the other wing 152, 154 may extend aft with respectto body 156, and the combustor dome 118 may be attached to the fuelnozzle assembly 90 or to another component.

Referring now to FIG. 5C, another exemplary embodiment of a separatecombustor dome 118 and outer and inner walls 102, 120 is illustrated. Inthe embodiment illustrated in FIG. 5C, the combustor dome 118 includes aforward-extending inner wing 154 but no outer wing 152; rather, an outerend 158 of the combustor dome 118 extends to the outer wall 102. Toretain the combustor dome 118 and seal against combustion gas leakagearound the dome, the inner wing 154 is pinned with the inner wall 120 atthe forward end 88 of the combustor 80, and the outer end 158 ispreloaded against the outer wall 102. More particularly, a springelement 160 is pinned with the outer wall 102 at the combustor forwardend 88, and the spring element 160 presses against the body 156 of thecombustor dome 118 to preload the outer end 158 of the combustor dome118 into a lip 162 defined in the outer wall 102. By utilizing themounting configuration illustrated in FIG. 5C, positive definiteretention and sealing of the combustor dome 118 may be provided whileminimizing thermal stresses in the dome, which is particularly usefulwhen the combustor dome 118 is made from a CMC material.

As previously stated, the outer wall 102, inner wall 120, and combustordome 118 may comprise a CMC material. More particularly, in exemplaryembodiments, the combustor portion 104 and the turbine portion 106 offlow path assembly 101 are integrally formed from a CMC material suchthat the resulting unitary structure is a CMC component. For example,where the combustor portion 104 includes the outer liner 108 of thecombustor 80 and the turbine portion 106 includes the outer band 110 ofthe first turbine stage nozzle portion 82N, the shroud 112 of the firstturbine stage blade portion 82B, the outer band 114 of the secondturbine stage nozzle portion 84N, and the shroud 116 of the secondturbine stage blade portion 84B, the outer liner 108, outer bands 110,114, and shrouds 114, 116 may be integrally formed from a CMC materialto produce a unitary CMC outer wall 102. As described above, in otherembodiments, additional CMC components may be integrally formed with theouter liner 108, outer bands 110, 114, and shrouds 114, 116 to constructa unitary CMC outer wall 102. Similarly, the inner wall 120 may beformed from a CMC material. For instance, where the inner wall 120comprises separate components, e.g., inner liner 122, inner bands 124,136, and blade platforms 132, each component of the inner wall 120 maybe formed from a CMC material. In embodiments in which two or morecomponents are integrated to form a unitary inner wall 120, thecomponents may be integrally formed from a CMC material to construct aunitary CMC inner wall 120.

Examples of CMC materials, and particularly SiC/Si—SiC (fiber/matrix)continuous fiber-reinforced ceramic composite (CFCC) materials andprocesses, are described in U.S. Pat. Nos. 5,015,540; 5,330,854;5,336,350; 5,628,938; 6,024,898; 6,258,737; 6,403,158; and 6,503,441,and U.S. Patent Application Publication No. 2004/0067316. Such processesgenerally entail the fabrication of CMCs using multiple pre-impregnated(prepreg) layers, e.g., the ply material may include prepreg materialconsisting of ceramic fibers, woven or braided ceramic fiber cloth, orstacked ceramic fiber tows that has been impregnated with matrixmaterial. In some embodiments, each prepreg layer is in the form of a“tape” comprising the desired ceramic fiber reinforcement material, oneor more precursors of the CMC matrix material, and organic resinbinders. Prepreg tapes can be formed by impregnating the reinforcementmaterial with a slurry that contains the ceramic precursor(s) andbinders. Preferred materials for the precursor will depend on theparticular composition desired for the ceramic matrix of the CMCcomponent, for example, SiC powder and/or one or more carbon-containingmaterials if the desired matrix material is SiC. Notablecarbon-containing materials include carbon black, phenolic resins, andfuranic resins, including furfuryl alcohol (C₄H₃OCH₂OH). Other typicalslurry ingredients include organic binders (for example, polyvinylbutyral (PVB)) that promote the flexibility of prepreg tapes, andsolvents for the binders (for example, toluene and/or methyl isobutylketone (MIBK)) that promote the fluidity of the slurry to enableimpregnation of the fiber reinforcement material. The slurry may furthercontain one or more particulate fillers intended to be present in theceramic matrix of the CMC component, for example, silicon and/or SiCpowders in the case of a Si—SiC matrix. Chopped fibers or whiskers orother materials also may be embedded within the matrix as previouslydescribed. Other compositions and processes for producing compositearticles, and more specifically, other slurry and prepreg tapecompositions, may be used as well, such as, e.g., the processes andcompositions described in U.S. Patent Application Publication No.2013/0157037.

The resulting prepreg tape may be laid-up with other tapes, such that aCMC component formed from the tape comprises multiple laminae, eachlamina derived from an individual prepreg tape. Each lamina contains aceramic fiber reinforcement material encased in a ceramic matrix formed,wholly or in part, by conversion of a ceramic matrix precursor, e.g.,during firing and densification cycles as described more fully below. Insome embodiments, the reinforcement material is in the form ofunidirectional arrays of tows, each tow containing continuous fibers orfilaments. Alternatives to unidirectional arrays of tows may be used aswell. Further, suitable fiber diameters, tow diameters, andcenter-to-center tow spacing will depend on the particular application,the thicknesses of the particular lamina and the tape from which it wasformed, and other factors. As described above, other prepreg materialsor non-prepreg materials may be used as well.

After laying up the tapes or plies to form a layup, the layup isdebulked and, if appropriate, cured while subjected to elevatedpressures and temperatures to produce a preform. The preform is thenheated (fired) in a vacuum or inert atmosphere to decompose the binders,remove the solvents, and convert the precursor to the desired ceramicmatrix material. Due to decomposition of the binders, the result is aporous CMC body that may undergo densification, e.g., melt infiltration(MI), to fill the porosity and yield the CMC component. Specificprocessing techniques and parameters for the above process will dependon the particular composition of the materials. For example, silicon CMCcomponents may be formed from fibrous material that is infiltrated withmolten silicon, e.g., through a process typically referred to as theSilcomp process. Another technique of manufacturing CMC components isthe method known as the slurry cast melt infiltration (MI) process. Inone method of manufacturing using the slurry cast MI method, CMCs areproduced by initially providing plies of balanced two-dimensional (2D)woven cloth comprising silicon carbide (SiC)-containing fibers, havingtwo weave directions at substantially 90° angles to each other, withsubstantially the same number of fibers running in both directions ofthe weave. The term “silicon carbide-containing fiber” refers to a fiberhaving a composition that includes silicon carbide, and preferably issubstantially silicon carbide. For instance, the fiber may have asilicon carbide core surrounded with carbon, or in the reverse, thefiber may have a carbon core surrounded by or encapsulated with siliconcarbide.

Other techniques for forming CMC components include polymer infiltrationand pyrolysis (PIP) and oxide/oxide processes. In PIP processes, siliconcarbide fiber preforms are infiltrated with a preceramic polymer, suchas polysilazane and then heat treated to form a SiC matrix. Inoxide/oxide processing, aluminum or alumino-silicate fibers may bepre-impregnated and then laminated into a preselected geometry.Components may also be fabricated from a carbon fiber reinforced siliconcarbide matrix (C/SiC) CMC. The C/SiC processing includes a carbonfibrous preform laid up on a tool in the preselected geometry. Asutilized in the slurry cast method for SiC/SiC, the tool is made up ofgraphite material. The fibrous preform is supported by the toolingduring a chemical vapor infiltration process at about 1200° C., wherebythe C/SiC CMC component is formed. In still other embodiments, 2D, 2.5D,and/or 3D preforms may be utilized in MI, CVI, PIP, or other processes.For example, cut layers of 2D woven fabrics may be stacked inalternating weave directions as described above, or filaments may bewound or braided and combined with 3D weaving, stitching, or needling toform 2.5D or 3D preforms having multiaxial fiber architectures. Otherways of forming 2.5D or 3D preforms, e.g., using other weaving orbraiding methods or utilizing 2D fabrics, may be used as well.

Thus, a variety of processes may be used to form a unitary structure,such as the outer wall 102 depicted in FIG. 3A, as a unitary CMCcomponent. More specifically, a plurality of plies of a CMC material maybe used to form each unitary structure. The plurality of plies may beinterspersed with one another to integrate the various portions formingthe unitary structure. As an example, the unitary outer wall 102 of FIG.3A may be made from a plurality of outer liner plies, a plurality offirst turbine stage outer band plies, a plurality of first turbine stageshroud plies, a plurality of second turbine stage outer band plies, anda plurality of second turbine stage shroud plies. Where the outer linerplies meet the first turbine stage outer band plies, ends of the outerliner plies may be alternated with ends of the outer band plies tointegrate the plies for forming the outer liner portion with the pliesfor forming the first turbine stage outer band portion of the unitaryouter wall 102. That is, any joints between the plies forming unitaryouter wall 102 may be formed by alternating plies on one side of thejoint with plies on the other side of the joint. As such, the plies forforming unitary outer wall 102 may be interspersed to integrate theplies and, thereby, each portion of the unitary outer wall 102. Ofcourse, the CMC plies may be laid up in other ways as well to form theunitary structure. In addition, laying up the plurality of CMC plies mayinclude defining features of the unitary structure or other component(e.g., inner liner 122 when not integrated with inner band 124 to from aunitary inner wall 120 or separate combustor dome 118 as shown in theembodiments of FIGS. 5A and 5B) such as openings 142 in combustorforward end 88, outer wall flange 144, and inner wall flange 146.

After the plurality of CMC plies are laid up to define a unitary CMCcomponent preform, the preform is cured to produce a single piece,unitary CMC component, which is then fired and subjected todensification, e.g., silicon melt-infiltration, to form a final unitaryCMC structure. Continuing with the above outer wall 102 example, theouter wall preform may be processed in an autoclave to produce a greenstate unitary outer wall 102. Then, the green state unitary outer wall102 may be placed in a furnace to burn out excess binders or the likeand then placed in a furnace with a piece or slab of silicon and firedto melt infiltrate the unitary outer wall 102 with at least silicon.More particularly, for unitary outer wall 102 formed from CMC plies ofprepreg tapes that are produced as described above, heating (i.e.,firing) the green state component in a vacuum or inert atmospheredecomposes the binders, removes the solvents, and converts the precursorto the desired ceramic matrix material. The decomposition of the bindersresults in a porous CMC body; the body may undergo densification, e.g.,melt infiltration (MI), to fill the porosity. In the foregoing examplewhere the green state unitary outer wall 102 is fired with silicon, theouter wall 102 undergoes silicon melt-infiltration. However,densification may be performed using any known densification techniqueincluding, but not limited to, Silcomp, melt infiltration (MI), chemicalvapor infiltration (CVI), polymer infiltration and pyrolysis (PIP), andoxide/oxide processes, and with any suitable materials including but notlimited to silicon. In one embodiment, densification and firing may beconducted in a vacuum furnace or an inert atmosphere having anestablished atmosphere at temperatures above 1200° C. to allow siliconor other appropriate material or combination of materials tomelt-infiltrate into the component. The densified CMC body hardens to afinal unitary CMC outer wall 102. In some embodiments, the final unitarystructure may be finish machined, e.g., to bring the structure withintolerance or to define openings 142 in forward end 88, and/or anenvironmental barrier coating (EBC) may be applied to the unitarystructure, e.g., to protect the unitary structure from the hotcombustion gases 66. It will be appreciated that other methods orprocesses of forming CMC components, such as unitary CMC outer wall 102,unitary CMC inner wall 120, or the like may be used as well.

Additionally or alternatively, other processes for producing unitarycomponents may be used to form unitary outer wall 102 and/or unitaryinner wall 120, and the unitary structure(s) may be formed from othermaterials. In some embodiments, an additive manufacturing process may beused to form unitary outer wall 102 and/or unitary inner wall 120. Forexample, an additive process such as Fused Deposition Modeling (FDM),Selective Laser Sintering (SLS), Stereolithography (SLA), Digital LightProcessing (DLP), Direct Metal Laser Sintering (DMLS), Laser Net ShapeManufacturing (LNSM), electron beam sintering or other known process maybe used to produce a unitary outer wall 102 and/or a unitary inner wall120. Generally, an additive process fabricates components usingthree-dimensional information, for example, a three-dimensional computermodel, of the component. The three-dimensional information is convertedinto a plurality of slices, each slice defining a cross section of thecomponent for a predetermined height of the slice. The component is then“built-up” slice by slice, or layer by layer, until finished. Superalloymetallic materials or other suitable materials may be used in anadditive process to form unitary outer wall 102 and/or a unitary innerwall 120. In other embodiments, a unitary outer wall 102 and/or unitaryinner wall 120 may be formed using a forging or casting process. Othersuitable processes or methods may be used as well.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. An outer wall of a flow path of a gas turbineengine, the outer wall comprising: a combustor portion extending througha combustion section or the gas turbine engine; and a turbine portionextending through at least a first turbine stage and a second turbinestage of a turbine section of the gas turbine engine, wherein thecombustor portion comprises an outer liner of a combustor of thecombustion section, wherein the turbine portion comprises an outer bandof a nozzle portion of the first turbine stage, a shroud of a bladeportion of the first turbine stage, an outer band of a nozzle portion ofthe second turbine stage, and a shroud of a blade portion of the secondturbine stage, wherein the combustor portion and the turbine portion areintegrally formed such that the combustor portion and the turbineportion are a single unitary structure, and wherein the single unitarystructure defines an outer boundary of the flow path.
 2. The outer wallof claim 1, wherein the combustor portion further comprises a combustordome extending across a forward end of the combustor and an inner linerof the combustor, and wherein the turbine portion further comprises aninner band of the nozzle portion of the first turbine stage.
 3. Theouter wall of claim 1, wherein the turbine portion further comprises aplurality of nozzle airfoils within the nozzle portion of the firstturbine stage.
 4. The outer wall of claim 1, wherein the outer band andthe shroud of the first turbine stage converge toward an axialcenterline of the gas turbine engine, and wherein the outer band and theshroud of the second turbine stage diverge away from the axialcenterline.
 5. The outer wall of claim 1, wherein the combustor portionand the turbine portion are integrally formed from a ceramic matrixcomposite material such that the single unitary structure is a ceramicmatrix composite component.
 6. A flow path assembly of a gas turbineengine, the flow path assembly comprising: a combustor dome positionedat a forward end of a combustor of a combustion section of the gasturbine engine; an outer wall having a combustor portion extending fromthe combustor dome through the combustion section and a turbine portionextending through at least a first turbine stage and a second turbinestage of a turbine section of the gas turbine engine; and an inner wallextending from the combustor dome through at least the combustionsection, wherein the combustor dome extends radially from the outer wallto the inner wall, wherein the turbine portion comprises an outer bandof a nozzle portion of the first turbine stage, a shroud of a bladeportion of the first turbine stage, an outer band of a nozzle portion ofthe second turbine stage, and a shroud of a blade portion of the secondturbine stage, and wherein the combustor dome, the combustor portion andthe turbine portion are integrally formed such that the combustor dome,the combustor portion and the turbine portion are a single unitarystructure.
 7. The flow path assembly of claim 6, wherein the singleunitary structure defines an annular flow path.
 8. The flow pathassembly of claim 6, wherein the combustor dome defines an opening forreceipt of a fuel nozzle assembly.
 9. The flow path assembly of claim 6,wherein the flow path assembly further comprises a plurality of nozzleairfoils extending from the inner wall to the outer wall within thenozzle portion of the first turbine stage, and wherein the plurality ofnozzle airfoils are integrally formed with the combustor dome, thecombustor portion and the turbine portion such that the plurality ofnozzle airfoils are a portion of the single unitary structure.
 10. Theflow path assembly of claim 6, wherein the single unitary structure isformed from a plurality of pre-impregnated ceramic matrix compositetapes.
 11. The flow path assembly of claim 6, wherein the single unitarystructure is formed from silicon-carbide fiber cloth.
 12. A gas turbineengine, comprising: a combustion section; and a turbine sectionincluding a first turbine stage positioned immediately downstream of thecombustion section and a second turbine stage positioned immediatelydownstream of the first turbine stage, wherein the combustion sectionand the turbine section define a flow path, wherein the combustionsection includes an outer liner defining an outer boundary of the flowpath through the combustion section, wherein each of the first turbinestage and the second turbine stage includes a nozzle portion and a bladeportion, each nozzle portion comprising an outer band defining an outerboundary of the flow path through the nozzle portion, each blade portioncomprising a shroud defining an outer boundary of the flow path throughthe blade portion, wherein the outer liner, the outer bands, and theshrouds are integrally formed such that the outer liner, the outerbands, and the shrouds are a single unitary structure.
 13. The gasturbine engine of claim 12, wherein the single unitary structure is aunitary outer wall of the flow path, the unitary outer wall defining aunitary outer boundary of the flow path through the combustion section,the first turbine stage, and the second turbine stage.
 14. The gasturbine engine of claim 12, wherein the combustion section furthercomprises an inner liner defining an inner boundary of the flow paththrough the combustion section, and wherein the inner liner isintegrally formed with the outer liner, the outer bands, and the shroudssuch that the inner liner is a portion of the single unitary structure.15. The gas turbine engine of claim 14, wherein the nozzle portion ofthe first turbine stage and the nozzle portion of the second turbinestage each comprises an inner band defining an inner boundary of theflow path therethrough, and wherein the inner band of the nozzle portionof the first turbine stage is integrally formed with the inner liner,the outer liner, the outer bands, and the shrouds such that the innerband of the nozzle portion of the first turbine stage is a portion ofthe single unitary structure.
 16. The gas turbine engine of claim 15,wherein the nozzle portion of the first turbine stage further comprisesa plurality of nozzle airfoils integrally formed with the inner liner,the inner band of the nozzle portion of the first turbine stage, theouter liner, the outer bands, and the shrouds such that the plurality ofnozzle airfoils are a portion of the single unitary structure.
 17. Thegas turbine engine of claim 16, wherein the inner liner, the inner bandof the nozzle portion of the first turbine stage, the outer liner, theouter bands, the shrouds, and the plurality of nozzle airfoils areintegrally formed from a ceramic matrix composite material such that thesingle unitary structure is a ceramic matrix composite component. 18.The gas turbine engine of claim 12, wherein the single unitary structureis formed from a plurality of ceramic matrix composite pliesinterspersed to form the single unitary structure.
 19. The gas turbineengine of claim 12, wherein the outer liner, the outer bands, and theshrouds are integrally formed from a ceramic matrix composite materialsuch that the single unitary structure is a ceramic matrix compositecomponent, and wherein the ceramic matrix composite component isdensified using a melt infiltration process.